text:
Why do many GA aircraft not have a "Both" fuel selector?
I've been trying to rack my brain around this and asked a few CFI's who couldn't give me more than "because that is how it was done", but a good number of GA aircraft save for Cessna and maybe a few others only have a Left/Right fuel selector:

Image Credit: Make Time For Flying Blogspot
To me, as an engineer, it makes great sense to have a Left/Right/Both selector such as found in most Cessna's. This lets the pilot focus on flying the aircraft rather than having to switch tanks every 30 minutes to balance fuel out. A Left/Right/Both selector isn't sigificantly more complicated than a Left/Right selector, so that leaves me wondering...
Why do GA manufacturers opt to use just a Left/Right selector valve given that a majority of GA accidents are fuel related and off-airport landings with fuel in a tank is still relatively common? Is there a safety issue I'm not seeing?
I understand that even with a "Both" selection, fuel may not drain from both tanks equally, but you can always switch to the fuller tank for a little while then back to both in that case.

Given that Cessna produced nearly 100k 150, 172, and 182 series aircraft compared to the roughly 33k PA-28 aircraft Piper produced, your title would better reflect reality if "most" were changed to "many" or similar.

@JonathanWalters When I say "most" I mean by model, not necessarily production numbers.

A quote from the book of regulations, Chapter 14, part 23, verse 951(b):

(b) Each fuel system must be arranged so that—
    (1) No fuel pump can draw fuel from more than one tank at a time; or
    (2) There are means to prevent introducing air into the system.

On a high wing aircraft satisfying #2 on that list is pretty easy: If you connect both tanks to the "top" of the fuel selector as in a Cessna 172 or similar then as long as you have liquid in at least one tank gravity will ensure there will be liquid at the top of the fuel selector to feed through into the engine:

The design of the system provides a passive "means to prevent introducing air" because gravity is always trying to pull fuel downward out of the wings and to the fuel selector.

On a low-wing aircraft like the Piper in your picture we don't have gravity to help us: There's no natural "head pressure" at the fuel selector (or at least much less than in a high-wing with the tanks mounted well above the valve), so we're relying on the fuel pump to draw fuel through the system.
This means we can no longer guarantee item #2 above: If the fuel selector were in the "Both" position and one tank ran dry the fuel pump is going to happily try to suck air through the line from the empty tank, and that could cause the engine to quit.
It's possible to guarantee item #2 in a low-wing fuel system using float valves that seal the tank outlets when the fuel is below a certain level - an active means to prevent introducing air into the system - but that adds complexity and weight.
Since the aircraft already needs to have a fuel selector valve anyway it's easier to just design the valve in such a way to satisfy item #1 on the list: Make sure the fuel pump can only draw from one tank at a time - i.e. "Don't provide a "Both" position on a low-wing aircraft."

I didn't think about the introduction of air, which is a good cause. I'd be willing to bet though that the reason for not using mechanical float valves to shut the empty tank off automatically may be more of "another point of failure" more than weight and complexity. Float valves have been around quite a long time and don't weigh more than a pound. I guess the other side is that it would introduce more unusable fuel (maybe a gallon at most) but I still would think that having the float valve is better than an embarrassing landing with fuel in one tank.

Another potential safety issue is if the selector was at "both" and one tank developed a leak. In that situation you don't want any means for the levels to equalize through the selector valve when it is set to "both" - especially if the purpose of the "both" setting is to reduce the amount of attention given to the fuel gauges!

Where Cessna does use a fuel pump, that pump is often designed to draw fuel from the header tank (for example in the C177), thereby also satisfying #1.

@JonathanWalters Indeed, IIRC the fuel-injected C172s are also designed this way (there's a "fuel reservoir" tank after the fuel selector, and before the electric boost pump), because in theory with the pump on and one tank dry you could draw fuel at a rate that exceeds the supply from the other tank (thereby introducing air). Depending on the size of the header tank this may satisfy the letter of the regulations while missing the spirit though: Running that header tank dry will still wake up the pilot in a hurry!

Good engineering would avoid the introduction of additional potential failure modes, so a method of isolating a nearly empty tank might not be such a good idea.

Another way of passively guaranteeing #2 would be to have the tank contain a collapsible fuel bladder that actually holds the fuel (think wine-in-a-box); as fuel is consumed, the bladder slowly collapses, avoiding air ingress into the fuel-containing part of the system.

Oh, and your link is broken.

On C177s, there is a plastic enclosure around the fuel tank selector.  If the selector knob is removed, the plate removed, and the knob replaced, then an Off position can be selected, which can be useful for maintenance. The plate makes sure that can't be selected normally.

Comment found on RPX Aviation about Left, Right, Both and OFF

If you carefully read the POH for early Cessna 172s, the POH cautions you about flying on both tanks when cruising at altitudes over 5000 ft.  Doing so may cause "power irregularities" (A.K.A. engine failure).  But I guess this is okay if it is in the POH.  Just switch to one tank, ignore the sputtering, glide for 60 (very long) seconds, then switch to the other tank and hope the vapor has cleared and the engine restarts.  That is why many early Cessna 172s have a placard on the fuel selector (between the seats, out of sight) that says something to the effect of "Use single tank above 5000 ft.

Another thought to add to the discussion. Some high wing aircraft are known for fuel to drain from one tank into the other when both tanks are selected. Airplanes seldom fly perfectly level all the time. This creates uncertainty for the pilot. He/she may keep a mental a note of how much gas they have, but incorrectly assume which tank it is in.  Yes the engineer could install check valves to prevent gas from unintentionally cross flowing back into a different tank, but that would require the introduction of check-valves which would add cost, weight and another failure mode.  KISS is a good approach. Best to provide a fuel system that is simple, reliable and allows the pilots to know for sure where their fuel is stored.

This is definitely true of the Cessna 182s. You could set the fuel selector on my 182 on one side and go flying and when you land you need to put fuel in both tanks. You also had to be careful when fueling if the plane was not level. Fuel will flow from one tank to the other even if the fuel selector is set to off.

@JScarry The Cessna fuel selector cross-feeds in the Off position just like in the Both position; it’s just missing the connection to the engine. In some models, the Off position was replaced with a separate fuel cut-off valve, so you could prevent cross-feeding.

@StephenS I just watched a Blancolirio video when a brand new C172 got stuck in CA because of the winter storms. The new ones have left, right, both, and a red knob above the fuel selector for fuel cutoff. https://youtu.be/ZlJ4d2pIAh4

@JScarry That’s what I was referring to.

Just a thought...in my experience aircraft without the "BOTH" selector position tend to have a higher wing dihede.  For example a Cessna 177 is about 1.15 degrees and a PA-28 is about 7 degrees.  So the BOTH might be used to address fuel unporting concerns.

Of course you have to take into account that low wing aircraft almost always have more dihedral than high wing aircraft of similar performance and handling, so we could state this alternatively as "low wing aircraft are more likely to lack a "both" position than high wing aircraft" --  not saying your answer is wrong, but just looking at it from a different angle  --

@quietflyer, you might re-read my comment.


text:
Why doesn't the flight path angle increase beyond a certain angle even if the angle of attack is increased?

From level flight, when the angle of attack is increased the flight path angle increases (the aircraft starts to climb). What causes the aircraft to stop to increase its angle of climb (angle between the horizon and the speed vector) as I continue increasing AoA at full thrust?
Is the effect described in Paragaraph 1 the same as with steepening the descent at idle thrust when I pull the stick?
Is it right that the critical AoA, above which continuing to pull the stick  causes decreasing speed vector's angle (i.e. "going down instead of up"), is the same for climb and descent and corresponds to $V_X$ (best angle of climb speed) up to negligible corrections?

Thanks
Alex

In three sentences or less, please state your question.

Sorry @Ryan it's stated in a very beginning, am I too imprecise wi this short formulation?
"In short (details below) the question sounds: what is the factor causing different behaviour, resulting in either increasing or reducing angle of climb, after the pilot pulls the stick back (raising AoA)."

@RyanMortensen thanks, I refined the question, thanks!

I'm not sure someone will take time to read your post, with the question is buried in the middle of a long post. My suggestion would be to create two identifiable sections in your post: one very crisp and synthetic with the question (s) stated very straightly. If this really adds something, a second section with your current ideas, but no additional request/question. The first section should be sufficient for the person able to provide an answer, the second would just help them to confirm the question is correctly understood. Just a suggestion.

Sorry, there is too much here to understand what exactly it is you are seeking. I think you need to remove at least 50% of the text to fit the user-friendly format of this site. In the meantime, if you have no excess power to climb then you are at the absolute ceiling, which is not really related to angle of attack. To help you, I think you should read about slow flight, induced drag, and power/thrust/drag curves.

I think I understand what you're asking.  The reason descent angle steepens with an increase in AOA at slow speeds is because the airplane is already below the best glide speed meaning that any reduction in speed (and increase in AOA) will cause the distance gained horizontally to decrease in relation to the distance descended.  It is because the induced drag is increasing, hingering the plane's forward/horizontal movement.  At least, that's the simplest explanation I can give you to assist you in understanding the concept although it may not be 100% scientifically accurate.

@RyanMortensen and others thanks for an attempt to read/understand and for your hints — I was imprecise and reformulated the question.

Are you increasing angle of attack, or are you increasing pitch angle?

@DeltaLima Since stick pulling increases AoA (page 2 from here http://www.ehfc.net/AngleOfAttack.pdf or 6.1.1 from here https://www.av8n.com/how/htm/aoastab.html#sec-aoa-stability), yes I'm increasing AoA, and change pitch then is composed of stabilized AoA + stabilized angle of climb and thus pitch is secondary to my question.

If you keep increasing your angle of attack, your drag will increase and combined with the increasing along trajectory component (your path angle will increase) of the weight vector will slow you down (assuming your thrust does not exceed your weight + drag). Because you slow down, your lift will decrease. The lift will now be less than the across trajectory component of the weight vector and the flight path will become less steep.
I don't understand your second point.
The critical angle of attack is where the maximum lift coefficient occurs. If you increase the angle of attack further, the wing will stall and the lift will decrease. This does not occur at Vx

Many thanks! It makes more clear for me!

I meant the same effect: when pulling stick decreases speed vector angle instead of increasing it. In normal speed descent (like 65 knots) when I pull lightly — I descend less steeply (please read carefully what I wrote: "angle between the horizon and the speed vector" goes up). But at very low speed descent when I pull the stick it starts to descent steeper (NOT stall). I was asking whether the cause is the same as in Paragraph 1.

With 3 again, I was not meaning going to stall (not "stalling AoA"). I was asking whether there is a special AoA (I called it "critical", sorry for a clumsy term) at which pulling stick will result not in speed vector going up, but going down. My guess it is $V_X$. I know that $V_X$ is higher than $V_{stall}$. Thanks for your help!

Thanks to all the commentators for pointing out partial answers.
I dare say I found the reference which answers why at the same power setting there exists a point (at some AoA) at which the vector of speed angle (= angle of climb/descent) starts reacting opposite to stick movements (pull stick => less climb angle / steeper descent).
It is power curve.

Power curve is best explained here from where I bring a citation (Paragraph 7.8):

The airplane is trimmed for a definite angle of attack, and hence a
  definite airspeed at 1 G. The yoke is part of the angle-of-attack
  control system. Pulling back on the yoke will always make you slow
  down.
If you are on the front side of the power curve and if you don’t mind
  airspeed excursions, you can use the yoke as a convenient, sneaky way
  to control altitude. This is because airspeed is linked to altitude
  via the law of the roller-coaster and via the power curve.
Warning: just because this works OK 99% of the time, don’t get the
  idea that it works all of the time. Bad habits are easy to learn and
  hard to unlearn. Do not get the idea that pulling back on the yoke
  always makes the airplane go up. On the back side of the power curve,
  it doesn’t work — and might kill you. In critical situations
  (including approach and departure), you simply must control the
  airspeed using the yoke and trim.

Power curve contains an implicit dependence of AoA (since AoA is roughly airspeed):

So when I am at full thrust climb and I start pulling, I increase AoA and move right to left on the power curve — increasing agle of climb. At some point for the reasons of fastly increasing drag coefficient and slowly increasing lift coefficient (see lift coeffcient and dreag coefficient curves) it stops increasing angle of climb. I'm in this point (which is $V_X$ by the way, see the link):

Exactly the same occurs in descent. In idle thrust regime, I start pulling stick back more and more and I move right to left on the power curve, descending less steeply.But after passing some speed (which is not $V_X$ but rather close $V_{L/D}$) I start descending steeper:

As explained in 1 and 2, the points on the power curve at which this happens are $V_X$ and $V_{L/D}$.

"Pulling back on the yoke will always make you slow down."  Unless your name is Patty Wagstaff.

@RyanMortensen not getting it even after reading about her. Explain pls?  Sorry for dumb questions, you know, I'm yet only a small beginner student trying to think over something before I convert it to the experience and the wisdom of flight :)

Sorry, that was meant to be a joke...  She is a world-famous aerobatic pilot, and if you're flying upside down and pull back on the yoke, you will speed up, not slow down.

@RyanMortensen :)) I thought of that but you know needed to check whether there MIGHT BE any other regimes where I miss something. Thanks for nice guidance through my first question here!

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text:
What options would Qantas Flights 7/8 have in case of emergency?

Route For Qantas 7&8
As this route shows, once the flight exits SYD, then all it has nearby is just the great pacific ocean. In such a situation, if it hits some emergency, what all options would it have, if it wants to land, due to some emergency like a patient on board?

as you can see from the map, the ocean is littered with islands, some (including Hawaii, about 2/5 of the way) have airports. If you want a flight that is really far from alternate airports, you want to look here: http://aviation.stackexchange.com/a/1347/1467

You should emphasize the medical emergency case, as almost all technical emergencies are taken into account in the certification process.

@ManuH "almost".  :)

Although it appears that going from Australia to North America, there is nothing but the great Pacific Ocean underneath. However the truth is (as Federico mentioned) that there are many airports along the way which can allow an A380 to land there, and many more close enough that an A380 can divert to in case of an emergency.
During the early days of this flight (Qantas 8 to be exact), it had to stop at Noumea, New Caledonia for a fuel stop as the pilots thought that it would be safer to do that.
Wikipedia has a list of diversion airports in Pacific. If you turn on the options to view nearby airports on FlightAware and zoom in on the map, you can find many airports.
It should also be noted that not all emergencies require an immediate landing.

There is a standard called ETOPS for two-engines aircraft on which they are certified. It specifies the maximum distance an aircraft can be away from the nearest airport while it flies.
For the A330, ETOPS is 180 minutes / 1,700 nm / 3148 km. It may sounds like a lot of flying in case of an emergency but the aircraft is designed with enough redundancy (2 engines / parallel and backup avionics system).
The standard is more important in case of a two-engines aircraft where losing one engine has more impact than on a four-engines aircraft.

authors: Юрий Волкович

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Is there any estimate of the number of cargo aircraft which can be handled and accommodated per hour at Chitral Airport in northern Pakistan?
I am working on a modeling case study in which I am using a realistic disaster response operation situation after a major earthquake in Chitral, Pakistan.
In this regard, I am using Chitral Airport as a local response center from where the transportation of relief to different affected areas will be realized.
I want to have an estimate of the number of cargo aircraft (such as C-130, C-17 or others in the inventory of neighbouring countries i.e. China and India) which can be handled on Chitral Airport in one hour.
So, the estimation quantity is total number of aeroplanes handled (landed and accommodated) per hour.

By looking at the said airport, one can pretty easily see that the limiting factor will be the offloading time of a cargo plane. The Apron is very small, and can only handle two, maybe three in maximum of cargo planes.
I'm not an expert on cargo handling, but a rough estimate would be about an an hour per plane for the C-17 size aircraft. A single operation would, with taxiing, take about 1,5 hours. Since the apron can only handle the aforementioned 2, max. 3 planes at a time, separation during flight would not be an issue.
So my best guesstimate would be a maximum of a bit less than two operations per hour for large aircraft, given the weather will not limit flying. Smaller planes could operate on much shorter interval, half and hour maybe. Then again, two large cargo planes would pretty much block the apron anyway, so adding smaller planes to the mix might not help much. edit: As the comments suggested the cargo was forwarded via air, this would complicate the equation. A single load of a C-130 would translate into about 4  Mi-17 loads which is manageable, but a fully laden  C-17 would require a whopping 15 Mi-17 flights (X. Managing 2 cargo plane takeoffs and landings and tens of helicopter operations on such a small area as the Chitral airport is not safe.
You should take into account what kind landing and t/o weights  Chitral can handle. If that turns out to be a limiting factor, there would not be much sense in operating large cargo planes half empty to/from there.
edit: Since some comments mentioned the unprepared field capability of C-130 and C-17, and it was also suggested that offloading chould be done outside the apron, I must stress that the unprepared field capability does not mean that the planes can taxi, take off and land everywhere. The load bearing capability of the surface used must be taken into account even if planes landing gears were designed to be able to operate on all kinds of surfaces! You do not want to bring the operation to a halt simply because a plane sunk its wheels to the ground. Once that happens, they need to be lifted out.
Also runway length can be a restriction. Even though (if I understood correctly) Chitral would be the destination of cargo, the landing distance requirements might be a limitation for heavy landings.
Taking all of the aforementioned aspects into consideration, large planes might not be the optimal choice after all. You should also think about the ground logistics. Running the ops at max efficiency will kick up considerable ground traffic.
X)Cargo capacities: C-130: 20000kg
C-17: 77000kg
Mi-17: 4 000 kg internal or 5000kg external

Actually we are not using land transportation, we are using helicopters for last mile delivery, so according to your suggestion, in 4 hours the airport should be able to handle 12-13 C-130 airplanes (full of humanitarian ration)

Keeping in mind that we are also using the airport for Mi-17 helicopter operations (although a cricket ground nearby can also be used)

The C-130 can use unprepared runways.  Can it also use unprepared aprons?  There are several fields next to the airport, but I cannot tell if they are flat.

According to US department of defense, C-130 can be operated using unprepared fields. Regarding surrounding fields, if you use the 3D view of the Google Earth and zoom in, as far as I can see, they are flat.

By the way, https://www.defense.gov/News/News-Stories/Article/Article/1648322/air-force-navy-officials-describe-need-for-c-130-aircraft/ is the link to the relevant DoD website

I think this answer is wrong since the aircraft would probably unload off the apron.

A C-17 could land on this runway, but it would need to be mostly empty.

Assuming you are sending aircraft which can unload on unprepared surfaces, and the flood plane to the east of the runway is a suitable surface:
The main limiting factors are

Safe separation.  This might be about 120 s between when the leading aircraft finishes taxiing and when the following aircraft lands.
Taxi time.  Assume the typical taxi is 700 m at 7 m/s.  Then add 100 s for taxi.

Unloading time and departure time are not relevant because you probably do not have enough cargo aircraft to fill up the space east of the runway (after demolishing small obstacles.)  Unloading can be simultaneous.  Just leave your empty aircraft there until no more are landing.
My estimate is about 16 aircraft per hour.  You will have a tough time finding 16 aircraft ready for this mission.

Does this answer assume as much taxiing, marshaling, and unloading on X acres of basically flat ground, as on the same area of paved tarmac? Sounds unrealistically optimistic. I've operated C-130's on unpaved surfaces for this sort of thing, but they were a far, far cry from a hastily prepared patch of flat ground. And even then, not nearly as efficient as operations on a paved ramp.

Is the estimate of 16 aircraft/hour based on that many C-130's? Or C-17's? Or 16 Cessna 206's? How long do you assume each aircraft takes to unload when reaching that estimate? How many aircraft are assumed to be unloaded at once to arrive at that estimate? What is the relationship between the numbers stated as the "main limiting factors" (i.e. 220 seconds) and the conclusion of 16 aircraft/hour? How can the statement "unloading time ... {is} not relevant" be justified? The difference between 48 hours to unload each aircraft, or 4 hours, or 40 minutes, or 4 minutes, seems highly relevant.

So how much C-130 do you suggest will be able to land on the said airport @RalphJ?

@Engr.MoizAhmad  Given the very limited data presented, I think Jpe61 gave you a very good answer, and I don't have anything much to add to it.

Great, thanks!!

@RalphJ 60*60/220 = 16.  I do not understand why that would need to be in the answer.

@RalphJ The question is about aircraft per hour.  I have modeled one hour.  After the first hour approximately, you have run out of aircraft to land.  If the question were about aircraft per week, unloading time would be very important.

@RalphJ Why are you asking about marshalling time?  This is a disaster scenario.  There would be no marshallers.

@AnonymousPhysicist aircraft operations are complex, a simple equation cannot take into account all considerations. And: should there be more than ten operations per hour on a single runway airport /-field, with many pilots unfamiliar with the place, you can bet you behind there will be marshallers there. Loads of them.

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How accurate are these Stall Diagrams depictions of relative wind?
Why in diagrams of a stall, the relative "wind" comes in horizontally no matter how the airplane is orientated? Doesn't the engine, which produces the thrust for the "wind", stay parallel to the wing making the diagrams of wind at higher AoAs much more inaccurate because of the direction of thrust?

Image source

Both diagrams are rather inaccurate. OK, the upper one is somewhat close to reality (only the streamlines are wrong), but the lower one is really lousy. In a superstall it is entirely possible for the fuselage to point nose-down.

@PeterKämpf, depending on how it is entered, deep stall can be both nose-up and nose-down; the important part is the angle of attack and that is correct. I agree behind the airfoils the streamlines don't look right.

The diagrams are accurate.¹
The thrust, produced by the engine, is one of the forces acting on the airframe. The others are lift, drag and weight (a.k.a. gravity).

(from How It Flies, chapter 4)
When these forces are in equilibrium, i.e. their (vector) sum is zero, the aircraft flies straight, in whatever direction it does (that is what Newton's first law says: an object remains at rest or continues to move at constant velocity unless acted upon by external force).
When the forces are not in equilibrium, the aircraft accelerates in the direction of their sum (that is what Newton's second law says: the sum of external forces on an object is equal its mass times its acceleration). But direction of acceleration and direction of velocity are independent. If the aircraft is turning, acceleration is to the side while velocity is (approximately) forward. If it is slowing down, acceleration is aft and velocity is still forward.
So thrust and direction of flight don't have to be parallel.
For a more practical argument, imagine an airplane doing the tailslide aerobatic manoeuvre. The nose points up and the engine is running and producing thrust up. But because the thrust is less then the gravity, the aircraft slows down until it stops midair and starts to fall back to the ground, tail first. So for a moment, the relative wind will come from aft while the engine is still producing forward thrust. Only then will the aircraft pitch over to restore normal airflow over the wings and return to fully controlled flight.
Now of course since the lift and drag forces depend on the relative wind. The lift, which is the aerodynamic force perpendicular to the relative wind, increases with angle of attack, which is the angle between the relative wind and the wing planform, until stall angle of attack is reached.
So when the plane is flying level slowly, the nose is pointed up to create sufficient angle of attack. The thrust is thus pointing slightly up, so small part of it balances a bit of weight, reducing the lift requirement. That is why aircraft actually stall at slightly slower speed when the engine is developing full power than when it is stopped.
Read the How It Flies chapter 4 for more details on the forces. And the rest of How It Flies; it explains the physics pretty well.

¹ That is accurate regarding the direction of oncoming relative wind. They are not accurate regarding the flow around and behind the wings.


text:
What is 'blending' for a turbine blade?
What is the process of 'blending' that is used as part of jet engine maintenance?

Blending is the process of filing and polishing damaged sections off of the turbine blade. This prevents further damage and extends the life of the engine without having to replace parts, making it a very economical solution. There is a limit however, as modifying the turbine blades leads to a small cost in efficiency that can build up over time.

Doesn't working on a blade affect the balance of the rest of the turbine? How is that restored? Must the entire assembly be re-balanced post-blending?

@curious_cat Good question. I think I'll post that as a separate question.

Howcome it's called 'blending'?

In that image, I'd assume that the post-blending blade is the one on the right. However, the one on the left actually looks like it's got more material on it. I wonder if it was a damaged blade that had additional metal welded on, then the blending is the grinding/polishing process necessary to get it back to the proper shape and to balance it with the rest of the fan.


text:
Why don't drones use ducted fans?
If efficiency of converting battery power to lift or thrust is paramount, why don't drones use shrouded or ducted fans, instead of unshrouded propellers?
This is the only image I could find to clarify what I mean by "ducted fan":

Image source (public domain)
My thoughts:

An extremely light weight duct could be made, so weight is probably not a factor
Increased suceptibility to winds could be a big negative
Maybe it doesn't make much difference at the low thrust levels of drones

Some do: http://www.horizonhobby.com/product/multirotor/multirotor-aircraft/bind-n-fly-15087--1/inductrix-bnf-blh8780

Don't underestimate the importance of weight; small quadcopter drones are meant to be quite rugged and survive a good tumble. A very light duct will be fragile, and anything heavier will deduct from flight time as more power is spent to remain in a hover. (that's why you also rarely see prop guards in use.)

Better late than never:  Here is a large hobby drone that uses ducted fans and channels exhaust through tubes:
https://www.youtube.com/watch?v=5L6FSdUmEpg&ab_channel=StefanoRivellini

The reason not many drones use ducted fans is efficiency.
Ducted fans are great for fast forward flight but not for static lift.
Explanation: Yes, the 'duct' reduces losses at the blade tips, However, adds significant losses at the intake lip and exit where the adjacent air is pulled into the flow. For ducted fans installed in aircraft, These losses are reduced at high forward velocities and increased at low forward velocity. For drones, there is situation with perpendicular intake flow, when drone is travelling sideways and fan pointing up, there is even greater losses at front intake lip.
The most efficient blade design for static lift is to have the blade as long as possible (like a helicopter) and travelling as slow as possible. Least efficient to have short blades travelling at high velocity.
It is easier to imagine what is going on if you increase the viscosity and imagine the fluid is water instead of air. Very large blades would move water only a small about, but over a large area. however, very tiny blades would be require to move vast amounts of water through a small area to produce the same thrust. moving that much water increases drag as the blades have to suck in water at the intake and the exiting water would be slowed very quickly by the surrounding water at the exit.

They are used - have been for almost 2 decades.  Check out Avid Aersospaces T-hawk.

AVID, in subcontract with Honeywell, assisted in the design of T-Hawk,
a ducted-fan micro air vehicle (MAV) as part of an accelerated DARPA
project. T-Hawk is an unmanned micro air vehicle that provides
real-time situational awareness in critical situations with over
30,000 hours total flight hours. (At the link)

The Ducted Fan has a narrow field of speeds where its efficiency is higher than an open propeller, or a turbine, see image, as elements in DF design, as distance of blade tip to duct and others, it was discussed in Aviation Stack Exchange, may change results, this is better tested experimentally, according to the desired use for machine.

Cons
Using ducts, weight will increase than advantage of getting smooth airflow to blades.
Using only propellers there will be more maneuverability    for sudden change in direction.
Pros
Using ducted fans, undisturbed airflow during windy or gust situations

The efficiency of the ducted fan can be increased by the shape of the duct. If the airflow is compressed, i.e. the outer airflow section is thinner of the inner section, the fluid is compressed (more compression means more speed in the airflow) and that's what gives most of the trust in normal propellers even if not ducted, if you note in small propellers airplanes, typically twin engines, there's a small porthole and another smaller one in the back of the engines, that's where the compressed airflow comes out, more or less similar design happens in jet engines I mean bigger airliners, not rockets. I was looking into building a ducted fan drone, but none of the commercially available ducted fans seems to follow this design.

Max, one writer actually tried to build what you described.  Forcing air through a smaller exit did not work,  for the mass flow crowd.  Propellers are airfoils and benefit, just like wings, by generating "lift" at lower angles of attack with MUCH less drag through the shape of the airfoil.

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/


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What are runway Aiming Point markers used for?
I'm interested in knowing what aiming points on runways are used for:

Why do some runways have aim points and some do not?
We say that pilots must use the maximum length of the runway. Therefore, why don't we touch down on or near the threshold lines as our aiming point?

Hi @adil. Have a look at this question.

This answer seems to adress most of it: http://aviation.stackexchange.com/questions/16714/what-determines-the-touchdown-zone-and-aiming-point-markings-on-a-runway?lq=1 Duplicate?

@curious_cat I agree that question is related, but this one differs at least in that it asks why some runways do not have aiming point markings

@curious_cat, I don't see that addressing the second point and that points seems the more important one here.

"We say that pilots must use the maximum length of the runway." Not sure where you've heard that, my goal in conducting landings is to use the least amount of runway possible without being hard on the brakes/tires. Usually this involves touching down as soon as possible, on the glideslope/airspeed. Maybe its just not worded the best? Often when doing short-field approaches you want to touch down as close to the threshold as possible, otherwise the touchdown point is just fine.

@RonBeyer To add to your list of requirements, we also don't want to accidentally land short of the runway either.

@Lnafziger Yes, good point, especially considering that with larger commercial aircraft, much of the airplane is behind you (like the A380, which is 239' long, meaning probably a good 225' of aircraft is behind you).

Aiming point and touch down points are different points. Aiming point is where you aim for before flaring. If I aim at the runway threshold and do the flare right, my Cessna 172 will touch down just ahead of the runway numbers. You can choose anything to be your aiming point. A touchdown at the threshold will need an aiming point on the field before the runway. The runway markers mainly provide reference for runway lengths remaining- typically 500 and 1000 feet multiples and sometimes runway midpoint.

I'll try to answer just the first portion. The aiming point marking basically serves as a visual aiming point for a landing aircraft.

Image from code7700.com
They are not found in all runways because the requirement is based on the aerodrome code number and they are not required in all cases. Aerodrome Reference Code from ICAO Annex 14 Volume I:

Table from ICAO Annex 14 Vol I
From the Same reference:

5.2.5.1 An aiming point marking shall be provided at each approach end of a paved instrument runway where the code number is 2, 3 or 4.
5.2.5.2 Recommendation.— An aiming point marking should be provided at each approach end of:
a) a paved non-instrument runway where the code number is 3 or 4;
b) a paved instrument runway where the code number is 1;
when additional conspicuity of the aiming point is desirable.

Something's up with that diagram... it shows two 500-foot distances adding up to less than a 1000-foot distance starting from the same point.

We say that pilots must use the maximum length of the runway. Therefore, why don't we touch down on or near the threshold lines as our aiming point?

No. Using maximum length of the runway just follows from more general rule to have as much margin of error as possible. But errors can happen in both directions. Landing long is more common, so it makes sense to leave the largest margin at the end of the runway, but landing short also happens, so some margin should be left at the start too. The 1000 ft is considered reasonable margin.
Note that for small aircraft, some advocate using the middle part of the runway and leaving about as much margin at both ends. However the PAPI and ILS can't be adjusted for each aircraft, so they are installed at the standard 1,000 ft past threshold mark.
Besides, pilots don't have to use full length of the runway. There is also LAHSO.

The markings on the runway are primarily required by the types of instrument approaches and the length of the runway. The aiming point marks at 1000 feet are required for precision approaches, they may be shifted a small distance to accommodate another crossing runway or for abnormal obstacle clearance on the glide slope.
Many precision approaches can have the plane pop out of the clouds at 200feet above ground so they don't have much time to adjust. If a bit misaligned they still need to be on a reasonable path to the actual runway and not the fence and field before the runway.
Remember that the same marks must be suited to many types of airplanes. 1000feet is not a long distance at 130knots in a jet but more than enough for a stop and go in a bush plane. Also to consider is that most airplanes require a longer takeoff distance than landing distance.


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What would a "Plus Sign" marking be on aircraft fuselage?
What is the plus sign found on some aircraft bodies?
The first time I saw this sign was on a Robinson R44 helicopter. The "plus sign" was located below the right front door.

got any pics? It's easier to know what you are talking about then.

Are you referring to a marking, something painted onto the aircraft? You don't mean something like this photo do you?

If it's what @JonathanWalters was referring to, but much smaller, it probably marks the location of a first aid kit.

I agree with @FreeMan, its most likely a first aid kit, although after looking through probably 100 R44 pictures on Google images, I didn't see a single one with this mark.

Yes, a picture or, at least, more detailed description is needed for this to be answerable. At the moment it could be anything from a first aid kit to indicating that the aircraft is a medical evacuation aircraft to indicating that the aircraft flies for Swissair.

@peyman do you have a photo, or can you give a more detailed description of what you saw? This is a great question, but we need more information before it can be answered.


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What are the advantages and disadvantages of mixing the bypass flow with the core flow?
Many Aircraft engines mix the bypass air with the core gases before exhausting to atmospheric pressure through a propelling nozzle. However, some do not mix and have separate nozzles for the flow of each. What are the advantages and disadvantages of each?
PS: I hope that I am clear now. Sorry if the question wasn't clear earlier.

You've got several very broad questions there. I'd suggest trying to narrow it down to one thing at a time so you can get a good answer, otherwise, this will probably get closed. It sounds like you're working on a design of your own - you might do better to define your current design (with pictures! or at least sketches), then list your alternate(s) for one area, and ask for pros & cons. When you get that sorted, ask the next question as a whole new question, referencing this one for background if necessary.

@FreeMan: I hope that now it is short.

It is short, but still unclear to me what you're after. Maybe it makes more sense to someone with more of an engineering background.

Don't let the close vote scare you off! Keep working at it to define your question better, and it will get reopened and answered.

In simple terms, mixing the exhaust flow (with any bypass ratio) is good for all the reasons except for construction and weight. It is aerodynamically more efficient; it reduces noise (exhaust noise makes most of the jet noise); it cools combustion chamber (and other downwind parts) more efficiently, etc.
However, making the mixing chamber on a high bypass / big engine would require enormous outer nacelle to cover the whole engine (and extend somewhat beyond), and the weight penalty and complexity/cost negates all the benefits.

Not sure I follow your argument, since this mixing is commonly done on most if not all "modern" high-bypass ratio engines. Granted, some of them are quite enormous (I'm looking at you Trent XWB with your 118" fan.)

@FreeMan, Trent XWB doesn't have a mixing chamber and has classical separate nozzles. Its nacelle is sizable, larger than it could have been in theory (perhaps because the fan pressure ratio is high, or for other aerodynamic reasons), but it would need to be some 50% longer to allow for steam mixing. Similarly, Trent 1000, GEnx, GE90 etc. are not mixing.

Advantage improved propulsion efficiency, stream Mn is affected by increase in temperature ie speed of sound changes.
Disadvantages longer nacelle required to mix cold and hot streams, this increases bypass scrubbing loss, nacelle drag and weight & cost.

Early jet engines were called turbojets where all flow went through the turbine. In other words there is no bypass. These are the cigar shaped engines you see in early airliners and military fighters. These types of engines are very loud and whiny and required high fuel flows to achieve the amount of thrust required.
Bypass allows a few things, it provides a cooling flow of air to mix with turbine exhaust and lower the overall exit temperature so you don't overheat your nozzle. The combination of the two exhausts can also see gains in thrust for given fuel flows. One of the most significant gains of bypass flow is noise reduction by having this outer flow "masking" the turbine noise. This is desirable in airliners.
There are two types of bypass:
Low bypass is where there is more turbine flow than bypass flow. This is common in business jets and military fighters with a ratio of around 1.5:1 or 2.0.
High bypass where there is more bypass flow than turbine flow. This is common in modern airlines with bypass ratios upward so 10.0:1.
Low bypass:

High Bypass:

In low bypass, the flow is remixed prior to exiting the nozzle, giving the cooler exit temperature typically to save the nozzle material. In a high bypass, the bypass air is allowed to flow outside of the engine after its high bypass compressor section, giving a very efficient airflow and adding to the overall thrust by moving a large volume of air outside the combustion and turbine section. This give greater fuel efficiency and much quieter engines for airplanes that can use them due to their size.
Below is an example of a low bypass F-100 engine which is found on F-15 and F-16 fighters.

This is the GEnx high bypass engine found on 747-8 and 787.

So what are the advantages for otherwise comparable engines? E.g. CFM-56 has separate streams and IAE V2500 has mixed stream, but have otherwise very similar parameters and are alternative engines for the A320-family aircraft.

just a nitpick, if the bypass ratio is over 1.0, then there is more bypass flow than core flow


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How wide is the cone of confusion above a VOR and NDB?
I saw in book that the cone of confusion of VOR is 40 degrees, and NDB is 50 degrees of each side! Isn't it too much?
Maybe it's a mistake, or old information(book printed in 2004)
Clarify please:
What is the real amount of cone of confusion of VOR and NDB?
What should I do when flew above VOR/NDB (in cones) in terms of navigation?

I think the term "cone of confusion" is much more apt.  When very near to a VOR/NDB, you may get confusing, incorrect answers, not "silence"

Cone of Silence was from the Get Smart TV series!

We do use 'cone of silence' within aviation - it typically refers to radar, though, not navaids.

Oxford book Radio Navigation 2 called either "cone of confusion" as well as "cone of silence".

The value +/-40° is correct, but need to be interpreted. The primary reason of the existence of a "silence" (confusion) cone is a design choice to increase navaid range. Range and "isotropy" are antagonistic.
However, when the aircraft approaches the confusion cone, say at 45° elevation, while flying at 9 km above the station, its slant distance to the station is about 13 km. Even with a small apparent power (say a few RF watts), the radio link being not attenuated by obstacles, is still very good. It should not prevent the signal from being demodulated.
There is a second phenomenon, which is more critical for a D-VOR: The Doppler shift is created by the antenna scan in the horizontal plane only, and is maximum when the receiver is on the station plane. As the relative elevation increases, the apparent displacement in the direction of the receiver is reduced. This reduction of the distance results in a reduced apparent velocity toward/away from the receiver, hence a smaller Doppler shift.

Without the correct shift, -480 Hz to +480 Hz, bearing determination is inaccurate (for a detailed explanation of VOR principles, see this answer). For conventional VOR there is a similar problem but related to the amplitude.

Detailed case: D-VOR
A VOR radiates a signal that can be interpreted in the horizontal plane, so the system is designed with antennas radiating horizontally. There are two types of VOR, the most modern is the D-VOR (Doppler VOR). A common antenna for D-VOR is the Alford loop antenna which is horizontally polarized:

Left: Source, right: Source
The D-VOR antenna system consists of a circular array of antennas:

Source
This array has a radiation pattern similar to this in the vertical plane:

How to interpret:

The straight lines are elevation angles, 90° being the zenith. The brown circles are a scale for power radiated, the large circle is the power measured in the direction where it is maximum. The next circle is where power is only -3dB of the maximum. Using decibels to measure ratios is frequent when working with radio signals because this is a logarithmic scale. -3dB is half the power, -6dB is 1/4 of the power, etc.

The red curve is the radiation power curve of the VOR for any vertical angle. It is maximum around 30° and minimum at 90°. For 60° its about -20dB which is a ratio of 1/100th only.

VOR cone of silence/confusion is where the signal is too small to be correctly interpreted by the VOR receiver, and you can see there is a strong signal fall around 50°-60° (that is 30°-40° from the VOR zenith)

As visible, power is mostly radiated to one side, the other side of the antenna array radiates only a parasitic power. This is due to the way the D-VOR works: Interfering pattern between all the antennas of the array. Power is destructed on the right side of the diagram, while it is cumulated on the left side.

The interfering pattern is rotated electronically around the vertical axis at 30 rounds per second (1,800 RPM, 30 Hz), so the antenna actually radiates power in an annular volume (donut).

Doppler cone of confusion
Specific to the D-VOR, due to the way this VOR works:

In the conventional VOR relative bearing is obtained by sensing the phase difference between a reference modulation (FM) and a varying modulation (AM) which depends on the direction the VOR antenna is currently pointing to.

In the Doppler VOR, the varying signal is actually obtained by electronically moving the center of the antenna array. This creates a frequency shift that is equivalent to the AM modulation of the C-VOR, but in frequency (and the D-VOR electronically "rotates" in the other direction so that the inversion between AM and FM modulation is not visible, ensuring compliance between C-VOR and D-VOR with the receiver).

However the Doppler effect can be sensed only when the receiver is on the same plane than the electronic "displacement", i.e. the horizontal plane. When the receiver is located too close to the VOR zenith, the frequency shift becomes small, and cannot be sensed correctly. This is why, even with an isotropic antenna, the system cannot work at high relative elevations. So there is actually no point to have an isotropic antenna for the D-VOR.

C-VOR and NDB
A conventional VOR (C-VOR) works with a vertical antenna which radiation pattern is naturally similar to the one above without the need for interferences. The antenna actually rotates physically.
NDB use long wire antennas of many kinds. The power is radiated with nearly no particular preference (omnidirectional antenna), but the radiation pattern is still similar to the C-VOR one, except that there is no need to rotate it. The donut volume is obtained with the antenna stationary.
As the NDB is omnidirectional, the energy in the horizontal plane is not focused as strongly as for the C-VOR and D-VOR. However NDB range is still reasonable because:

NDB uses the LF/MF portions of the radio spectrum which are less subject to propagation losses.

NDB receiver doesn't require a strong signal to locate the beacon, as simply detecting the carrier is in theory sufficient.

Why is that done this way?
You may ask why not using an antenna (or an antenna array) with a circular radiating pattern radiating the same power in any vertical direction. Such "isotropic" antennas exist in practical.
The reason is the same than for rotating the antenna: It's to concentrate power in one direction, to increase the range. In technical terms this is to increase the effective isotropic radiated power (EIRP).
A good similarity is the one of a light bulb and a torchlight:

A light bulb has an omnidirectional radiating pattern, it lights up in all directions, but the power received at a distant point is limited (unless the light bulb is a very powerful one).

A torchlight focuses light in a small solid angle, multiplying by a large factor the intensity in the privileged direction. A high intensity can be perceived in this direction without using a huge power, and the range is increased at equal power. However to scan all points in space, the torchlight must be rotated.

Isn't it too much?

No, it isn't. VOR's and NDB's are pretty old technology. VOR's were introduced in the 1930's and 40's, but are still commonly used today.
If you have a cone of confusion of 45 degrees to all sides, this means that, at 40.000 feet, the signals will be unreliable within about 6 miles from the beacon - which isn't really that much. At lower altitudes, say 3.000 feet, the cone of confusion will only be half a mile on all sides of the beacon. These inaccuracies are considered when designing procedures that depend on VOR's or other radio beacons.

I appreciate your use of real distance numbers in this answer, it's important to keep that in mind. One might see "45º cone of uncertainty" and think that makes for a very large volume, until reminded that aircraft travel much farther horizontally than they do vertically. As you point out, FL400 is not even 7 NM above the surface of the earth, while the standard service volume for a high-altitude VOR might guarantee reception out to 130 NM horizontally at that altitude.


text:
Where can I find aircraft design data?
I am looking for a library of different aircraft data, variables such as coefficient of lift, drag, etc.
I've gotten access to the POH/AFMS and the FAA TCDS' but they often modify the core data for operations and its not the true numbers, so to speak. I remember at university some textbooks had tables of the data but those can be difficult and expensive to access.
I would completely understand if this data is next to impossible to find as I can see manufactures would hold such info close to their chest.
I am specifically hunting for data on the C208 and C208B.

I get that the price of text books is obscene (it was when I was in college decades ago), but buying some used texts at the local college/university (particularly if they have an aviation program) or at an online book retailer (hmmm... where would one find one of those...) might be a reasonably inexpensive way of gathering this data if you really need it.

Resource location questions are not considered on-topic on this forum.

I once spent substantial effort digitizing the POH data for a C208 and working to fit a drag polar to the data.  This depends on your propeller model.
This kind of data is not generally published by the manufacturer.  The best you'll typically find is someone else's estimate.
Here is what I ended up with.
Sref = 279;   % ft^2
W = 8750;     % lbf
AR = 9.35;
oswalde = 0.86;
K = 1.0/(pi*oswalde*AR);

CD0 = 0.0348;
deltaCDpod = 0.0046;

polar = @(CL) CD0 + K*CL.^2.0;
polarpod = @(CL) CD0 + deltaCDpod + K*CL.^2.0;

I have used a slightly sanitized/rounded version of the above in numerous AIAA papers including AIAA 2015-1676.  This paper does not include details of how the polar was obtained, but it does provide a resource you can cite.

Some wings use NACA coded airfoils and indeed the 208 uses a blend of two wing root: NACA 23017.424, wing tip: NACA 23012. You can find info (and generate them) here. There is lift and drag info for the tip airfoil (23012) listed there but they dont seem to have the root foil listed at that particular side.
You can find NASA's test data for the 23012 airfoil here and some even older test data on it here.

That is what I am currently working with via XFLR5, but I am seeking the "official" or fully tested numbers.

@JeffreyMG see my edits


text:
Why is corrected engine speed is used for engines, even though the RPM of the engine is the same both on ground and at cruise?
I understand that mass flow rate has to be corrected because air gets thinner as altitude increase. But what I do not understand is why the corrected mass flow rate is always plotted with corrected speed. Does that mean for higher altitude, that the RPM sensor of the engine will give different readings compared to that in the ground at the same corrected mass flow rate? In short, what is the physical meaning of corrected engine speed?

Corrected speed allows performance characteristics of the jet engine to be easily compared across operating conditions, specifically across different ambient temperatures.
Let's just say that for a given engine, when ambient temperature is 59 degrees F, 90% physical N1 gives 50,000 pounds of thrust.
When ambient temperature is 100 degrees, the air in thinner.  So 90% physical N1 gives less than 50,000 pounds of thrust.  When ambient temperature is 0 degress F, the air is denser, so 90% physical N1 gives more than 50,000 pounds of thrust.
On the other hand, 90% corrected speed will always be 50,000 pounds of thrust, regardless of outside temperature.
So in our example, when ambient is 100F, corrected N1 will be lower than physical N1.  When ambient is 0F, corrected N1 will be higher than physical N1.


text:
747 Fuel Drain Ports
I read about the fuel systems on aircraft for a project I'm working on, and I was wondering about the "fuel drain ports" mentioned, which the document stated were used for complete drainage of fuel tanks and sampling. On a 747-300, do they exist? What do they look like, where are they, and how are they controlled or opened/closed?

Disclaimer: The B747-300 belongs to the B747-100/-200/-300/SP family. I have detailed knowledge of the B747-400 & -8 which differ quite a bit from the originals. That being said I would be very surprised if Boeing changed much regarding sump drains.
All tanks in transport category aircraft have individual sump drains. That is a requirement in FAR 25.971. In the B747 the sump drains are fully mechanical poppet valves. The poppet is held in place by a spring pressing down on it sealing it against the valve seat. To drain it you press against the poppet unseating it and therefore allowing fuel to flow.
There are two versions:

direct type (fuel intake directly above the valve)
Used in no. 2 & 3 main tanks, the center wing tank and all surge tanks. Located under the wing in the lowest locations of said tanks.

indirect type (fuel intake connected to the valve via a fuel sump drain line)
Used in no. 1 & 4 main tanks and in no. 2 & 3 reserve tanks. Located under the wing inboard of said tanks. That's due the lowest point of the tank being above the engine strut and the associated fairing.

Cross-cut of a direct type sump drain:

I can’t help but wonder what the draining device looks like, and what the draining interval is.  I am picturing someone out there with a ladder and a GATS jar every morning.

@Max for the draining interval, see here: How often is the fuel sampled in turbine aircraft?

@Max the draining device is a transparent container with a long tube attached to it so you can easily reach the drains without a ladder.

Whew! I was thinking about the poppet drain for the oil filter in my car. I could only envision the poor sap who had to do this getting soaked in Jet A every time... :)


text:
Are there any companies that use atom interferometers in aircraft?
Are there any companies that use atom interferometers in aircraft? If so, are they used for navigation? If not, is it possible to assemble it?
Here is the link from wiki Atom interferometer and link from MIT Technology Review of it and a link from a company work on it.

A link to the description of the principle of AI would be useful.

Could you add a quick mention of what an atom interferometer is to your question, for people who haven't heard of them?

One company is working on AI.

Probably not.
The 2012 paper Advances in Atomic Gyroscopes: A View from Inertial Navigation Applications* states in the abstract that "there are still lots of problems that need to be overcome to meet the requirements of inertial navigation systems." Even once atomic inferometer gyroscopes are able to meet the basic requirements of inertial navigation systems, it will take years of testing and certification before they will actually be used in production aircraft.
I highly suggest reading that paper if atomic physics is your thing, but for us the conclusion is gold:

Recent progress on atomic gyroscopes has demonstrated that AIG is an
  ultra-high precision gyroscope for strategic grade inertial navigation
  in the future, and ASG is a high performance gyroscope which features
  both high precision and compact size. There are still lots of problems
  that need to be overcome when considering the atomic gyroscope for
  inertial navigation applications, however with the technologies
  available in the near future, the AIG could be used in space for
  fundamental physics research, the ASG based on comagnetometer can be
  used in strategic grade gimbaled INS, and the navigation grade ASG
  based on NMR can be demonstrated on chip-scale size. Furthermore, with
  the development of new theories and technologies for atomic
  manipulation, the future of high performance inertial navigation
  applications using AIG and ASG is indeed bright.

I would be slightly more cynical about time scales than this paper. This intro to INS notes on page 5 that it took 18 years for ring laser systems to reach maturity. No current atomic gyroscope system can meet the performance of ring laser systems, and the highest-performance AIG systems are on the order of a cubic meter. Chip-scale ASG systems have been produced, but their performance is dreadful compared to mechanical or ring laser systems.
*Note: I'm not sure how much to trust this article. It's published in a peer-reviewed journal, but there are clear copyediting errors. I don't have a better source, though, so I'll trust it for the purposes of this question. Also, note that this paper does not only have aviation users in mind; there are a lot of references to strategic INS, which I think refers to submarines, and it also talks about INS in the context of satellites.

I have glimpsed the paper you linked. AIG has ultra-high precision and ASG is compactable and sensitivity, which one is more suitable for aircraft up to date? Or which one could be application easily and faster?


text:
What is meant by the sentence, "It relates to aircraft handling speeds and circling area development."?
What exactly is meant by the sentence:

It relates to aircraft handling speeds and circling area development.

For which values does the phrase "aircraft handling speeds and circling area development" in the description for the PANS-OPS label indicate that PANS-OPS documents and rules are used? Because there is a label on the top left, such as "Standard", "Std/State", which indicates from which source the minimum and visibility values were created. What exactly is the difference between them?
Also, should I assume that the PANS-OPS rules are used in the formation of the circle-to-land values in the minimum table and the aircraft categories (A, B, C, D) based on their velocities?

Suggest searching for "PAN-OPS vs. TERPS". You'll find several presentations describing the differences in detail.

"Handling speeds" relates to the speed range airplanes can use when maneuvering around the airport to land on a given runway in the circling configuration (usually gear down and partial flap), as given in the 4 categories, and "circling area development" relates to the radius of the turning circle that results when maneuvering at a given max speed.
The max knots for each category relates to the ability to circle and keep the airport environment in sight and maintain obstacle clearance at that speed, allowing for the radius of turn at a standard turn rate at the maximum speed for the category.
The Category you use is based on the slowest speed you are able to use in the landing configuration at Max Landing Weight, and you use the category whose max speed is at or more than your max gross weight's reference speed, if a reference speed for that airplane is published (if no reference speeds are provided, you have to use 1.3 VSO at max landing weight). You want to use the slowest category you can fit into, since that gives the lowest minima.
For example, in the CRJ, the Vref at max landing weight ranges from 120-141 kt, from the airplane's Speed Book. So based on that you could use Cat C minima right up to the Max Landing Weight of 47000 lb where you just squeezed in (In the CRJ you actually circle at 160kt, the minimum speed for Flaps 30, and you go to landing flap as you got established on a base leg for the target runway and ready to commit, and slow to Vref plus a little margin extra for turning allowance on the turn to final).


text:
What is this hole at the bottom of the DA40NG body?
This is the bottom of the fuselage of a DA40NG aircraft and where the right wing attaches to the fuselage. What is the thing I'm pointing at with the red arrow in the picture?

As it is written in the photo, we see the fuel cooler inlet and outlet sections. But towards the right, where the wing is very close to joining the fuselage, there is a hole and a metal protrusion coming out of it.
Does anyone know what it is and what it does? At least from which source can I learn this information? Because I could not find this information in AFM and Airplane Maintenance Manual.

The fuel cooler is odd. It must be for certification using mogas or something.

That is a fuel sump drain valve. Water is heavier than fuel, so it sinks to the bottom of the tank. The sump drain valve is at the bottom of the tank so any water in the tank can be drained using the valve, this is done using a cup with a stick pointing out of it - you push the valve open using the stick and fuel and/or water drains out into the cup for visual inspection.
Airplanes often have multiple sumps, sometimes more than one per wing, either because they have multiple tanks, or multiple low points in a single tank. IIRC (and I could be wrong as it's been awhile since I flew one), the DA40 has 5: 2 per wing and one under the engine compartment. Most light singles have 3.
A pilot should drain fuel out of the sumps as part of the pre-flight checks to make sure there's no water contamination.

So, how is this fuel sump drain valve different from the fuel tank drain section on the left? Because before the flight, we check whether there is sediment or water by taking a fuel sample from the fuel tank drain section.

There's more than one @pilot162, I'll edit with more detail.

@pilot162 The 172 I fly has 6 per wing and 2 on the belly. Makes checking for contamination...tedious

We use only 1 drain hole for each fuel tank. I just learned that others do the same.

6 drains in a 172? I've never seen more than 1 per wing on a 172, I guess that means they were made before the lawyers took over @Jamiec!

@GdD Nor had I before this one. Its craziness

Been trying to find you a good picture, If you look super close you can just make out 4 of them https://www.flickr.com/photos/egbj/50394831848/

The number of fuel sump valves is dependent on the model of the Cessna. There is a difference between a 172E, 172G, 172M, 172N, and 172SP. Those are the only ones with which I am personally familiar.

The Cardinal has 4, one for each wet wing tank, one for a little 1qt collector tank in the belly, and the gascolator drain.

Schematic illustration of Drain Valve:

That is the bolt for the landing gear. You would open the panel with the "fuel cooling vents" and you can unscrew the bolt to take out the landing gear.

Your answer could be improved with additional supporting information. Please [edit] to add further details, such as citations or documentation, so that others can confirm that your answer is correct. You can find more information on how to write good answers in the help center.

Why do you think it is the bolt for the landing gear? I don't see any sources that say that. Please provide any supporting information.

Also, if you're unsure of something that you think might answer OP's question, you should probably post as a comment.


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What are different components of an engine pylon?
What are the different structural components that are inside the fairing of engine pylon and what are the functions of each part?
The pylons for each aircraft are made of Ti-64 alloys, can they be made using Aluminium alloys or the temperature requirements are very high so Aluminium cannot be used?
Thanks

This FAA document on "lessons learned" regarding fuse pins gives a lot of detail about the structure of the 747 pylon and its shortcomings.

Your question is very broad.  Are you looking for something specific? There isn't really a single configuration where one can simply list standard function of each component, but  I can give broad overview.
Firstly, pylon is also a conduit for multiple systems between aircraft and engine and thus contains secondary structure for mounting them. It looks much like aircraft wing with frames and ribs to support distributed loads. Secondary structure is easily made from aluminum and includes brackets, cleats, supports, stand-offs, and attachment hardware like clamps, nuts, bolts. Also pylon skins/fairings can (depending on design) be minor part of load distribution. But I presume you mean primary structure components that are most directly associated with mounting  the engine.
The functions are to attach/support engine (including withstanding emergency/crash/FBO loads), transmit thrust to airframe, and accommodate engine thermal expansion. Most generically, this is achieved by structural members forming a torque box or space frame. But realization and component names vary a lot. The are two fundamental families depending on attachment: 1) fuselage (i.e.) side-mounted,  and 2) wing (i.e. vertically)-mounted. We won't go into exotics like above-wing & wing-root mounting. Furthermore, there are few styles of wing-mounting (with different components) that fundamentally are driven by relative location of engine to the wing and relative size of engine to aircraft. There are also different mounting philosophies like soft-mounted (with elastomeric vibration isolators) or hard-mounted (without such).
Bear in mind that aircraft structures are designed to distribute load across multiple members and provide redundant load paths, so as to intentionally avoid single point of failure. Having said that,  4 key  elements can be identified:

a forward mounting beam/strut, which is perpendicular to engine axis and usually terminates in a yoke/frame that wraps around engine fan case. It primarily supports engine weight, but also crash loads & lateral FBO loads, including subsequent windmilling.
a thrust strut/link/mount, which is most aligned with engine thrust line, but inevitably runs diagonally to connect engine to pylon structure. It carries most of the thrust load into airframe, but would also carry longitudinal crash loads.
An aft/rear mounting beam/strut, which is also perpendicular to engine axis. Depending on engine size & proportion, in can vary from a frame to yoke to very thin struts. Aft mount's (primary but not only) function is to provide 2nd support point to secure the engine and allow expansion.
Space frame further into the body that disperses the loads and provides rigidity out of principal plane created by elements 1-3.

In a fuse-mounted configuration the principal plane is horizontal and functions are more decoupled. In wing-mounted configurations the principal plane is vertical and some functions & load distributions are more coupled.  Other designs like B-747, where engine hangs far below wing attachment points, rely on a space fame of axial members to suspend the engine in a 3D truss (as can be seen in previous post with link to FAA accident investigation). The functions mentioned before are even more coupled & shared between all members. The names are really irrelevant as Boeing chooses to refer to them as upper/ lower links and side braces.

In terms of actual parts, this is realized by main fittings, sleeves, bushings, alignment pins, load-bearing pins/bolts, etc.
Materials: Actually aluminum is widely used in mounts, with steel as required, and titanium when $/kg weight trade studies justify it. Titanium is expensive & horrendously difficult to machine and is used sparingly, albeit more so on recent designs that push margins. Of course, as everything in such highly complex products, it depends on specifics and sizing of a given design. Temperature must be considered, but is typically not a driving factor, especially for fuselage- mounted  engines,  and areas adjacent to cold sections of the engine. Actually it works out quite well - the heaviest and most loaded section of engine (like fan)is in front, which is cold. The loads are best supported & reacted by a big, intricate mount with large x-section area & moment of inertia (like yoke). Manufacturing it from titanium would be cost prohibitive and it is best made from aluminum. The aft hot section of the engine is supported by simpler, smaller struts that could be made from titanium. For wing-mounted though, the whole pylon area is subject to heat soaking, but weight is more evenly distributed between forward and aft mounts, so having smaller, simpler titanium fittings is practical.

Thanks for the information. That helped a lot.
I wanted to ask specifically about the main load carrying truss like component of the pylon. It's mainly made with Titanium alloys, but I was wondering if it could be made with Aluminium alloys or are there temperature constraints that should be considered?

I'll post my answer separately as there is character limit to comments.

Continuation of answer to clarifying comment.
I see, so your question was really about element #4 - the space frame that continues to carry loads further into main structure like wing box. To my knowledge, there are really no definite distinguishable sub-elements with universal names. They may be called beams (at least the main ones that hang engine off on any cantilever-like member), stringers, braces, bulkheads/frames - typically depending on shape & orientation to aircraft axes. Moreover, depending on  design's simplicity/complexity and ability to refine structure further, they can vary from few straight-forward main members  to many dozens of links.
I think it's best to get functional idea of load distribution by referring to a patent here:

https://patents.patsnap.com/v/US9889942-aircraft-assembly-comprising-a-mounting-pylon-primary-structure-integrated-to-the-structure-of-the-wing-element.html
For example of extreme (theoretical) refinement, research "pylon topological optimization".
For example of  real-life, analyzable design search examples on the web, like these pics from case study here

https://code7700.com/case_study_american_airlines_191.htm
I am sure that once you realize the fundamental load picture, you will be able to identify those elements by function, regardless of particular name.
> if it could be made with Aluminium alloys or are there temperature constraints that should be considered?

For such complex machines that require thousands of hours to design & cost  millions of $, the answer is - absolutely YES, temperature constraints are considered. The  question is - can we still get away with aluminum or are we forced to go to titanium. The answer depends on many factors & is  calculated for particular aircraft. To summarize broadly - it depends on config, how hot the engine runs, and what is heat load imposed by systems.  It helps to understand factors driving material selection and historic perspective.  As everyone is chasing fuel efficiency, the engines are running hotter for increased thermal efficiency. At the same time the regs have gotten stricter on ignition prevention, forcing lower permitted bleed temperatures near fuel tanks. This requires using higher capacity intercoolers with greater heat rejection than ever before. And they are inevitably located in the pylon. Moreover, there is trend to More Electric Aircraft (MLA) with bigger generators, and power transmission generates a lot of heat as well.
So if you have smaller, cooler, less efficient engine, which typically goes with smaller planes with less consumption load and tail/side mounting - you might be able to get away with aluminum. However, with a bigger, hotter-running engines, with greater bleed/electrical consumption load of modern airliners, and with under-wing mounting, leaving the pylon exposed to rising heat - you pretty much can't escape using titanium.

You should add this to / edit this into your existing answer instead of creating a new one (unless you reached the maximum character limit of 30,000).

At least this gives me a chance to give you two upvotes for what ist essentially one answer.

Understood, I considered it, but thought it would not not make sense in terms of flow with clarifying comment appearing after. Thanks for note.


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Did these parts come from a T-33 crash?
A Lockheed T-33 crashed near my hometown (Outer Cove, Newfoundland) in 1956. This year we have had some very strong storm surges, and a piece of metal has washed up on the shore. I am posting this picture to see if anyone  can tell if it is part of the crashed T-33. Thanks.

Am I right to assume the oval hole is approx. 2x8 inches?

Yes that is correct the are two on that piece.

oh hey, you found that old piece of dental work I lost when I was in NF last year...

@CiderJoe , awesome, do you know why they believe at the time it crashed?  IT would be amazing if you could find a bullet hole, fuel residue or whatever is relevant to the event!  Awesome idea on the museum.

They think it was weather. Here's the story on it if your interested.

https://www.pressreader.com/canada/the-telegram-st-johns/20150505/281500749811940

@CiderJoe -- did you ever get in touch w/ the guy who posted this answer on 4-6-2021, now deleted?  "I've also found two pieces of a aircraft from Outer Cove Beach within the last two days. I have pictures if you would like to see. I can email them."

It looks like the nose skin panel with the gun ports that some T-33s had.  So I'd be willing to put serious money down on "yes".

Lockheed T-33 Shooting Star ‘69330 / TR-330’ (N651)
Alan Wilson from Stilton, Peterborough, Cambs, UK

Damn you John, just as I was about to post an answer I noticed you were (way) faster. I totally agree with you, the top side is a bit different, but the rivet hole pattern on oval thingy are a match. So my money is on yes too.

Yeah a lot of posts are just races to Google Images aren't they lol.

It might be good to editorialize this answer a bit with what the questioner should do about it -- is there someone who should be notified? Can they legally keep it as a souvenir or sell it to someone on eBay?

How come the lower right corner of the plate is cut so neatly in a diagonal line?

Wow that looks like it for sure. The plane slid about 40 feet and then slid down over a 200 foot clift. Thats

We have a museum that has a section on the plane crash so I will be donating it to them.

@GlennWillen Why? That wasn't part of the question.

@Mast It depends on local jurisdiction, but generally speaking, debris from a crashed plane is highly sought after by authorities and in some cases it may be illegal to possess, or you may be legally required to report it to the NTSB or whatever agency is responsible for aviation accidents in your jurisdiction. https://en.wikipedia.org/wiki/Aviation_archaeology explains more. This is especially important when it pertains to crashed military aircraft, as explained in the article.

@zovits that looks like a fold in the metal.  You can see a rivet hole right at the fold.

@Nzall Considering this is about Canada, asking the TSB for advice won't hurt, although I'm not sure those care about military craft this old.

Seriously you guys?  It's a little piece of sheet metal from 1956!  The investigation closed its books on this one over 60 years ago, I doubt very much that there is a clue in this that any investigator could use to solve some lingering mystery about root cause.  There are far more compelling and recent crash sites with debris still out there that people no longer care about.  I would keep it as a souvenir or donate it with a clear conscience and not pay any attention to the naysayers.

@MichaelHall Calling people that recommend getting a solid legal statement on handling military debris naysayers ... you are the absolute mad lad!

@GlennWillen, one of the good things about *.stack is that it focuses on providing answers to questions asked, and is not filled with a lot of editorializing.  That being said, every user can edit or suggest edits to every answer.  If you feel you can improve the answer, you should do so.  Other users will provide feedback they feel appropriate.

Legal issues aside, do archeology, investigators, and the museum a solid and document when and where you found it. Without that it’s just a hunk of metal.

@candied_orange I think that some manufacturers have geotagging enabled by default, which the museum could use if one sends the original file.

I don't see any reason at all to editorialize this fine answer, I'm afraid @GlennWillen.  Such things are appropriate for comments.  aviation is (easily) the most sensible, consistent, and least-idiotic of all the SO sites.  (People who couldn't care a hoot about airplanes, read this site for that reason alone!!)  Pls do not let standards slip.

I am a retired journalist and lifelong aviation buff/writer/researcher in St. John's. I conducted an in-depth research project into this crash and published the results in a two part, front page story in the Telegram newspaper in 2015. Since the crash of a T-33 is the only one known to have taken place in Logy Bay, I'd say the chances of it being a piece of the plane are excellent (it is obviously a piece of aircraft aluminum).


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Why are there no large turboprop passenger airliners?
If turboprop engines are more efficient than turbofan/jet engines at lower speeds, and have a lower specific fuel consumption, why are there no large passenger airliners that utilize the turboprop engine?
I expect this would increase flight times, but say for example an airline was to only operate short-medium haul flights within Europe, wouldn't it be in the interest of minimum fuel burn to use the more efficient engine, the turboprop?
Most modern airliners seem to use two turbofan engines. I am curious as to why the largest turboprop aircraft have a passenger capacity of below 100 (ATR 72 and DHC Dash 8) and why there are no 150+ pax turboprop airliners.
With minimum fuel burn as the priority, ignoring costs and flight times, would this be a desirable and achievable design option?

Welcome to Aviation.SE! We already have similar questions such as Why are propeller engines uncommon on airliners? Let us know if you're looking for further details not answered there.

Possible dupe of Why aren't there any widebody propliners for the high-density short-haul markets? (disclosure: asked by me).

Large airliners don't fly at slower speeds.
It's simple economics.
If you have to connect two destinations 10,000 miles apart, you'll do maybe 2 round trips a week. Your passengers have no other choice, except a flight that takes even longer. You also have the luxury of flying between two very large cities.
If you have to connect two destinations 100 miles apart, you can do 3+ round trips a day. Your passengers choose and pick among a dozen other carriers that do the same, and many will just go by car or bus. And at this distance, at least one of the cities being services is small.
As a result, a single 50-passenger turboprop running a local route can transport as many passengers as a 500-passenger ultra-long-range turbofan. At the same time, the local route will have far fewer passengers per day.
Fast long-range turboprops are possible (see the Tu-95 and -114), and only a little slower than turbofans. But they're still louder. All in all, for large planes on long trips, both the flying public and people on the ground choose quieter faster planes over somewhat more fuel-efficient ones.
Extra noise for an hour is less of a pain, speed matters much less, and the smaller turboprops working these routes are less of a pain on the ground too. The fuel economy gap isn't all that large.


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Turbine blade manufacturing/assembly process
I want to know the process for manufacturing turbine blades and the companies that are involved.
Who makes the turbine blades? Who applies the thermal barrier coating? Who repairs the blades?
Let's take the General Electric GEnx as an example. Does GE make the blades themselves, or do they buy them from a supplier? Where do companies like MTU Aero Engines come in?

You'll have a hard time finding an exact answer, as that's unlikely to be public information, but here are some general trends:
In the case of the high pressure turbine, the blades are one of the key company jewels.  These are one of the highest technology pieces of the engine, and contributes a large part of the competitive advantage of one engine over another. Thus, this technology tends to be kept as a trade secret, and the parts are generally manufactured by the engine OEM themselves, and not outsourced.
For example, let's just say that tomorrow P&W discovered some new material that would increase EGT margin by 50 degrees.  That would give them a huge advantage over GE.  So they would want to keep it as closely guarded as possible.  If GE ever found out how to make that material, the competitive advantage is gone. So they want to keep all of the information as closely guarded as possible.  What's the best way to do that?  You make everything yourself.  If you outsource the manufacturing, then your material is out there in a place where you don't have control over it.  One or two blade might "walk away", and suddenly show up somewhere else.  You don't want that to happen.  So you keep it all in house.  In fact, you probably keep the information on a "need to know" basis.  I.e. the engineers who design the compressor or the fan aren't told the details of this new material.  They know it's there, and maybe some general information, but the details of how to make it aren't even widely known within the company.
Now that's for new engines. Ones that are actively being sold, where there is a competition, and customers are willing to pay a premium for performance.  Now consider on the other hand a very old engine.  Something like a CFM56-2, or a JT8D. That's ancient technology, so there's not really any competition on performance.  But there are still a lot them flying around, and they do need spare parts from time to time. The key driver for the spares market is going to be cost.  No one will pay a premium for performance here.  So you might see some of the manufacturing outsourced in this case.  You'll even see 3rd parties selling aftermarket parts for these engines (look up "PMA").

This makes a lot of sense. Thank you. I just have one small question. MTU said that they are actively participating in turbine blades in GE9x. how are they contributing to it?
Also, since they manufacture their own blades do they also do the TBC themselves? Then what is the role of companies like IHI, MTU?

Too many acronyms...

I can't say anything specific about MTUs participation in turbine blades as I don't have any specific knowledge. I would be surprised if they are making HPT blades but I don't know for sure. Maybe they are making LPT blades? That's still an advanced area, but not as much as HPT.


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What are the differences between a compressor wash on wing vs on test bed
I work in an aircraft engine overhaul facility and I’m trying to understand the process of compressor washing engines on Wing compared to washing them in a test bed. What process steps are different on wing? Are there time/noise restrictions on wing? The cost of washing an engine on wing? Who completes the comp washes?

You may transform the title in a question as this is a Q&A website


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What is the difference between position and navigation lights on light training aircraft?
What is the difference between position and navigation lights on light training aircraft (or general)? When I search for both, I find the same lights: red on the left, green on the right, and white on the back... Are they the same thing or is there a difference?


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What is the bulge under the engine of the Mirage III and some of its variants?
While attempting to create a Kfir in 3D modelling software, I noticed a bulge under the engine on the reference picture (a 2D side view) I was using, which I initially assumed to be a fin. After looking at some 3D images of it, however, it is far too wide to be a fin:

There is a similar bulge in the same position on the Mirage III:

The Atlas Cheetah and Mirage 5 both have a bulge in that position as well, although every non-Mirage-III-based Mirage seems to lack it. The Mirage 2000 has a bulge in a similar location; however, it appears to be much smaller:

My initial thoughts were either ECM or some kind of fuel tank. Does anyone else know what it might be?

Can't speak of the Mirages, but on the Kfir it's a fuel tank plus flares dispenser (the red squares) and RWR antenna (the black square). I wouldn't rule out a fuel tank on the Mirage as well, but on the 2000 it appears to be too small to hold any significant amount of fuel

On the Atlas Cheetah C, D, and E, it was a fuel tank and housed the downward firing chaff/flare dispensers.
[Underside]:

It was known as the "rocket bay" tank. The original Mirage III had a rocket fitted there, for maximum climb performance to intercept the ruskies coming over from the east.
Later models, including the Aussie Mirage III O, had the rocket system removed and replaced by a fuel tank. Wikipedia refers. If you look closely, you'll see the outline of the rocket bay: rectangular-shaped, with the faired piece extending rearwards. The tank could be removed.
It looks as though the Kfir has got a equipment pod in there.
I don't know what the small hump on the M2000 is.

Is there any source to support the claim of a "rocket bay"?

OP did not refer to the JATO mounting locations. https://pbs.twimg.com/media/BkNQohlCIAEkml4.jpg

@DeepSpace this picture seems to support this answer.


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What is the highest indicated airspeed ever achieved in a traditional aircraft?
It is fairly well known that the SR-71 holds the record for highest speed for a traditional aircraft (not a rocket/able to land and takeoff under its own power). However since the SR-71 flies much higher than other aircraft its indicated airspeed is comparatively low.
I was wondering what the fastest indicated airspeed achieved is and by which aircraft?

The highest would probably be when Space Shuttle lands into the atmosphere deep enough to be called an aircraft.

Touch I would not call the Space Shuttle a "traditional aircraft"...

@tsg A flying (falling) brick with winglets is still an aircraft I guess.

There have been a few aircraft that could get well above speed of sound at sea level.

I've found this: https://www.airspacemag.com/daily-planet/sultan-speed-180970800/, which seams to indicate that the low altitude speed record is just under 1000 mph.

define "traditional aircraft". By my definition the SR-71 would classify, by yours apparently not?

Sorry If I was unclear. The SR-71 would classify by my definition, but since it flies so high its indicated air speed is low (maximum of 400 knots).

By my definition, an SR-71 is not traditional because it is a hybrid between a turbojet and a ramjet with afterburner. What is incredible about the modified J58 is that it had more to give. Some say Mach 5. Limiting factor was heat of the airframe.

The world low-altitude speed record was set by Darryl Greenamyer in a rebuilt F-104 on 24 October 1977 at 988.26 mph (869.67 kts; 1'590.45 km/hr) and it wasn't really at sea level (mud lake Nevada on a hot day density altitude can be quite high so we don't know KIAS exactly). That airplane is by most considered to be the fastest due to small wind and frontal area. And it might also become one of the fastest land vehicle ever as some clever guy thought that getting rid of the wings was a good idea...

The SR-71 had a designed airspeed limit of 500KEAS. It was design for high altitude flight and they did not bother about structural strength to withstand low altitude fast flight or high g loads flight as you can see in the full flight manual.
Other jet fighter were known to fly fast at low altitude during mission like the F-4 with operating procedures of 1.34 Mach (863 mph, 759.5kts, 1,389 km/h).
Finally more recently the F-16, not known as a fast airplane was nonetheless designed the limit of 800KIAS in mind. The canopy was meant to melt at 850KIAS and the engine supposedly would blow up at 854KIAS. And numerous reports show that the pilot had to pull back the throttle in order not to exceed the 800kts limit at sea level.
The reason why we don't have higher speed recorded is just that there are no need for it, the aircraft where therefore not designed to exceed "reasonable" speed and no-one was willing to push the limit further. The risk versus benefits to bet on some engineers over-designing their aircraft so that they would fly above their limit speed is just not worth it. But I'm pretty sure if you where to "floor" a MiG-29/31 or a f15 up to failure you would certainly exceed those speed. It's like the longest flight, some records won't probably ever be tried again.

Re: "It's like the longest flight, some records won't probably never be tried again."  Yep, just like sailplane endurance records, all you prove is that the wind can blow against a ridge for a really long time, and a pilot can stay awake for a really long time, until he doesn't.  So no one keeps track of those records any more.

Yeah but the 64days straight in a Cessna will remain my absolute favourite/supidest one ^^

"won't probably never"?

Typo, feel free to correct them by editing the answer if you find any more of them.

Practically all supersonic jets are speed-limited by structural strength rather than by thrust at low (and mid) altitudes. MiG-29 and 31 are too different to be written like 29/31; perhaps 25/31 would do, but like SR-71, they were designed for high Mach rather than high IAS. MiG-29, on the other hand, is officially limited to 1500 km/h, which is 810 KIAS.

@Zeus correct, but we all know Russian engineering is all about stronger things. They still have rockets from the 60's flying perfectly well. So maybe we could excess the 810 KIAS design limit on the MiG29 as the structure might be a bit over-designed, and I added the MiG31 as even if it was meant to intercept SR-71, it had such oversized engines thrust wouldn't be the issue and the structure looks more sturdy than a SR71


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Which compressor is more like to surge and stall - axial or centrifugal?
I am studying about stall and surge in rotodynamic compressors and want to understand which one is more prone to these phenomena.

Axial for sure.  Its blades are wings subject to aerodynamic stall within the blade array.
Centrifugal compressors just throw air outward; they don't really "stall" if in-flow is restricted, but will surge due to the pressure drop and resulting load reduction on the compressor.  Centrifugal compressor surge is what is happening when you block a vacuum cleaner and it goes vwwwwwweeeeeeeeeeeeeeeeeeeeee.
Axial compressors, being an array of little spinning wings, depend on a controlled angle of attack on the blades and are much more sensitive to flow disruptions.  And once they do stall, the surge and flow disruptions throughout the engine are more severe (those whomp whomp sounds and flames coming out the front of the engine as reverse pressure waves are induced, that sort of thing).
So if you want efficiency and high compression ratios, you go with axial and live with the sensitivity, the high cost and complexity.  If you want insensitivity to inflow disturbances and can live with a larger diameter engine and less efficiency, and want to make the engine cheaper, you go with centrifugal.

Blade stall can occur in centrifugal compressors, depending on the difference in whirl velocity. Not in a vacuum cleaner, no.

Well they call it stall as in flow disruption but I was trying to differentiate between that and stalling of the airfoil blades of axials being closer to a regular wing.  I used the vacuum cleaner analogy because when air cycle machines get inflow restrictions (from letting the heat exchangers get dirty) they act like vacuum cleaners and surge and overspeed, and in the case of ACM spools they usually proceed to destroy themselves.

Why is axial more prone to stall? If the pressure difference over at least part of the compressor increases above what it can produce, either will stall. The difference is in how much the pressure has to drop for them to recover, and how quickly they'll recover, but the condition to cause it is basically the same.

Well yes but the centrifugal can tolerate larger pressure disruptions before stall/surge sets in.  On the 1st engines like the axial Jumo vs centrifugal Whittle, where the crude fuel metering system consisted of more or less a needle valve connected to the throttle with the pilot's eyes and brain the fuel control system, the Jumo had to be far more carefully managed, mostly because it was axial.


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How does landing gear affect thrust on a light twin-engine airplane?
I'm working on my commercial multiengine rating.  And I got this question from my examiner, and I answered: landing gear produces penalty drag, so it decreases thrust.
 But the examiner says this isn't correct... ?

Thrust is not decreased.  The drag penalty means that more thrust is required to achieve a performance condition than without the drag penalty.  In level flight, more thrust is required to go the same speed.  When climbing, more thrust is required to achieve the same climb rate.  Your maximum thrust available is what the engine can produce, so the extra drag subtracts from whatever surplus thrust is available.
In a light twin the key thing is the surplus thrust margin available for climb at the most efficient speed for climbing while single engine (blue line speed).  To fly in level flight single engine at blue line speed, we'll say you need 500 lbs of thrust from the good engine.  To climb at that speed requires more than 500lbs, and to get a decent rate of climb, say 400 fpm, you might need 800 lbs (about what you'll get with a Lyc O-360 running flat out), which we'll say is what the engine makes at full power at sea level (a piston twin that can climb better than 300 fpm single engine at gross is a pretty good performer).
Add more drag by leaving the gear down, and the thrust required just for level flight goes up from 500 lbs to, say, 700 lbs because of the 200 lbs of drag of the gear sticking out.  Instead of 300 lbs of surplus thrust margin available for climb, you only have 100 lbs and it's a struggle to get any climb at all or to avoid descending, and only if you use perfect technique (on-speed, 5 degrees bank into the live engine).
So the proper answer for leaving the gear down is, surplus thrust available for climb is decreased and therefore climb performance at maximum thrust is decreased.  Which in a light twin at gross on a hot day means, as the old saying goes, "the remaining engine takes you directly to the scene of the crash".